It is known that a gyrodyne, also called a gyroscopic actuator and often denoted by the acronym CMG (Control Moment Gyro), is distinguished from reaction wheels, commonly used for controlling the attitude of a satellite by exchange of angular momenta, in that the control moment gyro includes a rotor driven (by a motor) so as to rotate about a rotation shaft which is itself fastened to a support, called a gimbal, which is steerable (by at least one other motor) about at least one gimbal shaft fixed relative to the platform of the satellite, the axis of rotation of the rotor moving perpendicular to the gimbal shaft, whereas a reaction wheel is driven (by a motor) so as to rotate at a variable speed about an axis of rotation that is fixed relative to the platform of the satellite.
Agile satellite attitude control methods and systems of the prior art generally comprise a cluster of three or four control moment gyros delivering large torques along the three axes of the satellite.
One also well-known method consists in using two head-to-tail control moment gyros (their angular momenta being equal in modulus and opposed in direction) for producing torques in a direction, in this case the bisector of said angular momenta.
Moreover, patents U.S. Pat. No. 5,681,012 and U.S. Pat. No. 6,360,996 describe a method using two control moment gyros to produce torques along two different axes.
For this purpose, and with reference to FIG. 1, which shows schematically the arrangement of the two control moment gyros by the orientation of their gimbal axes and angular momentum vectors developed relative to the reference orthogonal coordinate system (X, Y, Z), the gimbal axes A1 and A2 of the two control moment gyros are mounted in the plane defined by the two axes X and Y of the coordinate system, this (X,Y) plane being orthogonal to the Z axis, which is for example the pointing axis of an instrument on board the satellite and which is intended to be tilted. The angle φ between the two gimbal axes A1 and A2 must necessarily be nonzero in order to obtain the desired effect. According to the two aforementioned US patents, the preferred angle φ is 120° C. The angular momenta H1 and H2 of the two control moment gyros are thus constrained to move in the planes P1 and P2 respectively, these being orthogonal to A1 and A2 respectively, and making between them the same angle φ. In the canonical position, the angular momenta H1can and H2can of the two control moment gyros are advantageously aligned in a head-to-tail configuration along the Z axis, so that the total angular momentum of the pair of control moment gyros is zero. This arrangement is called a “skewed scissor pair”
Starting from this canonical configuration, the angular momenta H1 and H2 of the control moment gyros are each pivoted about their respective gimbal axis A1 or A2 in such a way that the resultant torque has nominally a zero component along the Z axis, without which at least a third actuator, acting along the Z axis, would have to compensate for this component, which could be high owing to the fact that the torques delivered by the control moment gyros are very high.
In order for this component along the Z axis to be zero, it is necessary to constrain the temporal movement of the rotation angles L1 and L2, given to the two control moment gyros respectively, about their respective gimbal axis A1 and A2, from the canonical position.
More precisely, according to U.S. Pat. No. 5,681,012, it is necessary that:
                              ⅆ          L                ⁢                                  ⁢        1                    ⅆ        t              ·          sin      ⁡              (                  L          ⁢                                          ⁢          1                )              =                              ⅆ          L                ⁢                                  ⁢        2                    ⅆ        t              ·          sin      ⁡              (                  L          ⁢                                          ⁢          2                )            that is to say, by integrating:cos(L1)=cos(L2)+constant, the constant being zero since L1=L2=0 at time t=0.
Consequently, in order for the control method according to U.S. Pat. No. 5,681,012 to be able to be implemented, it is essential that the rotation angles of the control moment gyros, from their canonical position, be equal in absolute value, it being possible for the angles to have the same sign (L1=L2) or opposite signs (L1=L2). The skewing of the two gimbal axes A1 and A2 with a nonzero angle φ then ensures generation of torques in two different directions U1 and U2 in the (X,Y) plane, depending on whether the signs of said rotation angles are the same or are opposed, as described in detail in U.S. Pat. No. 5,681,012, to which the reader may advantageously refer for further details about this subject.
However, it is important to note that, in principle, the generation of these two torques can be accomplished, according to this known method, only sequentially and not simultaneously, as it is not possible to have L1=L2 and L1=L2 at the same time.
The first consequence of this known system and known method is the noncontrollability along the three axes of the system for small angles. Other actuators must be used to overcome this drawback. In addition, to tilt the Z axis about any axis U in the (X,Y) plane, it is necessary to decompose the rotation R(U) about the U axis into a product of two rotations, the first of which takes place about the U1 axis (R(U1)) and the second about the U2 axis (R(U2)).
Thus, to generate the rotation R(U), the satellite will firstly be tilted along U1 in order to perform the rotation R(U1), then along U2 in order to perform the rotation R(U2), with a stop phase between the two rotations.
The limitations of this method are therefore noncontrollability at small angles and also considerable suboptimization in the performance of maneuvers at large angles.
Patent U.S. Pat. No. 6,360,996 relates to improvements made to the method according to U.S. Pat. No. 5,681,012. The basic principle, namely the skewed scissor pair configuration, is maintained. However, in addition, deviations with respect to the constraints:
                              ⅆ          L                ⁢                                  ⁢        1                    ⅆ        t              ·          sin      ⁡              (                  L          ⁢                                          ⁢          1                )              =                              ⅆ          L                ⁢                                  ⁢        2                    ⅆ        t              ·          sin      ⁡              (                  L          ⁢                                          ⁢          2                )            that is to say L1=L2 or L1-L2, are accepted in U.S. Pat. No. 6,360,996, the disturbing torques induced along the Z axis then being compensated for by a variation in the speed of the control moment gyro rotors. Thus, complex couplings appear between the control along the (X,Y) axes and the control along the Z axis, in particular in maneuvering mode.
These couplings are not easily manageable and they induce the risk of saturation of the actuators along the Z axis. Management of this saturation is a central feature of the method, as results from the description given in U.S. Pat. No. 6,360,996, the more so as the control method described in that patent uses only very conventional tilt guidance concepts, by determination of trajectories and generation of torques to be applied to the satellite in order to perform the determined trajectories.
To alleviate the aforementioned drawbacks of the prior art (use of two control moment gyros to create torques along an axis, or along two axes, but with strong implementation constraints), a satellite attitude control system is disclosed that comprises a pair of control moment gyros and at least a third actuator in a configuration different from those known from the prior art, in particular the patents U.S. Pat. No. 5,681,012 and U.S. Pat. No. 6,360,996, so as to achieve attitude control along three axes of the satellite, and also rapid tilts, with guidance and control laws that are very simple to implement, and with controlled inter-axis couplings.
For this purposes the method according to the invention, for controlling the attitude of a satellite equipped with an attitude control system in a reference coordinate system (X, Y, Z) for positioning the satellite, and comprising at least three actuators called main actuators, two of which are control moment gyros each having a rotor driven so as to rotate about a fixed rotation axis with respect to a steerable gimbal that can be oriented about a gimbal axis perpendicular to the rotation axis of the corresponding rotor, and stationary with respect to the satellite, is characterized in that:                the gimbal axes of the two control moment gyros are fixed so that these gimbal axes are parallel to each other and to the Z axis, the angular momentum vectors ( H1, H2) of the control moment gyros therefore moving in the (X,Y) plane and making between them an angle (α) which, by definition, corresponds to a skew ε=180−α between the angular momentum vectors ( H1, H2) when α is different from 0° and 180°;        in addition to the two control moment gyros, at least a third main actuator is used as a complement, delivering torques in both senses in at least one direction not lying in the (X,Y) plane, so that this third main actuator is called the Z-axis main actuator;        a nonzero skew angle (ε) between the angular momentum vectors ( H11, H2) of the control moment gyros is imparted, said skew angle (ε) preferably being chosen to be small enough not to create an excessively large internal angular momentum on board the satellite but large enough to ensure controllability of the attitude control system along the three axes (X, Y, Z) without necessarily having to modify the rotation speed of the rotor of at least one of the control moment gyros;        the kinematic and dynamic variables, which are necessary for controlling the attitude of the satellite, such as for example the attitude angles and angular velocities of the satellite along the three axes, are estimated from measurements provided by sensors used on board the satellite;        setpoint variables, intended to allow objectives assigned to the satellite attitude control system to be achieved, such as for example the tilting and pointing along at least one of the three axes of the (X, Y, Z) coordinate system, are calculated; and        control commands are calculated, from differences between said estimated variables and said setpoint variables, and then sent to the main actuators, these control commands being intended to control the change in said differences over time, said control commands transmitted to the control moment gyros comprising at least commands intended to vary the orientation of their gimbal axes, such as for example gimbal angular position setpoints that have to be generated by a local position feedback control, or electric current setpoints, for currents that have to be injected into motors for orienting the gimbal axes, etc.        
This method using one pair of control moment gyros in this particular configuration, in which the angular momenta change in the (X,Y) plane with a nonzero angle α, about a position not aligned head-to-tail but with a nonzero skew angle ε=180−α, and also at least one third actuator for creating nonzero torques about the Z axis normal to the (X,Y) plane, is advantageous over the prior art in that it makes it possible, as described below, on the one hand, to very simply control the attitude of the satellite along the three axes (X, Y, Z) without it being necessary to modify the rotation speed of the control moment gyro rotors and, on the other hand, to easily perform rapid tilting maneuvers of the Z axis, by advantageously applying the guidance techniques in maneuvering mode that are proposed in the Applicant's patent FER 2 786 283, all this with great ease of design of the control system, in particular with regard to management of the coupling between the (X, Y, Z) axes and the design of the actuators that result therefrom.